Engine assembly for an auxiliary power unit

ABSTRACT

An auxiliary power unit having internal combustion engine(s) in driving engagement with an engine shaft, a generator having a generator shaft drivingly engaged to the engine shaft, a compressor having an outlet in communication with the internal combustion engine inlet, and a turbine having an inlet in communication with the internal combustion engine outlet. The turbine may be a first stage turbine, and the assembly may include a second stage turbine having an inlet in communication with the first stage turbine outlet. A method of providing electrical power to an aircraft is also discussed.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. application Ser. No.15/616,187 filed Jun. 7, 2017, which is a continuation of U.S.application Ser. No. 14/750,207 filed Jun. 25, 2015, the entire contentsof both of which are incorporated by reference herein

TECHNICAL FIELD

The application related generally to engine assemblies and, moreparticularly, to such engine assemblies used as auxiliary power units inaircraft.

BACKGROUND OF THE ART

Aircraft auxiliary power units (APU) commonly provide pressurized airand controlled speed shaft power to the aircraft systems as analternative to extracting this energy from the main engine compressorflow and accessory gearboxes. The APU is often used to power systemswhen the main engines are shut down.

Known ground-based APUs typically include added weight and complexitywhich may not be compatible with use in aircraft applications.

SUMMARY

In one aspect, there is provided an auxiliary power unit (APU) for anaircraft, the APU comprising: at least one internal combustion engine indriving engagement with an engine shaft; a generator having a generatorshaft in driving engagement with the engine shaft to provide electricalpower for the aircraft, a compressor having an outlet in fluidcommunication with an inlet of the at least one internal combustionengine; and a turbine having an inlet in fluid communication with anoutlet of the at least one internal combustion engine.

In another aspect, there is provided an engine assembly for use as anauxiliary power unit for an aircraft, the engine assembly comprising: atleast one internal combustion engine in driving engagement with anengine shaft; a generator having a generator shaft in driving engagementwith the engine shaft to provide electrical power for the aircraft; acompressor having an outlet in fluid communication with an inlet of theat least one internal combustion engine; a first stage turbine having aninlet in fluid communication with an outlet of the at least one internalcombustion engine; and a second stage turbine having an inlet in fluidcommunication with an outlet of the first stage turbine; wherein atleast one of the turbines is in driving engagement with the compressor.

In a further aspect, there is provided a method of providing electricalpower to an aircraft, the method comprising: flowing compressed air froman outlet of a compressor to an inlet of at least one internalcombustion engine of an engine assembly; rotating an engine shaft withthe at least one internal combustion engine; driving a generatorproviding electrical power to the aircraft with the engine shaft througha drive engagement between a shaft of the generator and the engineshaft; and driving a turbine of the engine assembly with exhaust fromthe at least one internal combustion engine.

DESCRIPTION OF THE DRAWINGS

Reference is now made to the accompanying figures in which:

FIG. 1 is a schematic view of a compound engine assembly in accordancewith a particular embodiment;

FIG. 2 is a cross-sectional view of a Wankel engine which can be used ina compound engine assembly such as shown in FIG. 1, in accordance with aparticular embodiment;

FIGS. 3-5 are schematic views of flow distribution assemblies which canbe used in a compound engine assembly such as shown in FIG. 1, inaccordance with particular embodiments;

FIGS. 6-7 are schematic views of cooling assemblies which can be used ina compound engine assembly such as shown in FIG. 1, in accordance withparticular embodiments;

FIG. 8 is a schematic view of a compound engine assembly in accordancewith a particular embodiment;

FIG. 9 is a schematic view of a compound engine assembly in accordancewith another particular embodiment, which may be used with the flowdistribution assemblies of FIGS. 3-5 and/or the cooling assemblies ofFIGS. 6-7;

FIG. 10 is a schematic view of an engine assembly in accordance withanother particular embodiment, which may be used with the flowdistribution assemblies of FIGS. 3-5 and/or the cooling assemblies ofFIGS. 6-7; and

FIG. 11 is a schematic view of a gas turbine engine in accordance with aparticular embodiment.

DETAILED DESCRIPTION

Referring to FIG. 1, a compound engine assembly 10 is schematicallyshown. The compound engine assembly 10 is particularly, although noexclusively, suitable for use as an airborne auxiliary power unit (APU).The compound engine assembly 10 includes an engine core 12 having anengine shaft 16 driving a load, shown here as a generator, for exampleto provide electrical power to an aircraft. Other possible loads mayinclude, but are not limited to, a drive shaft, accessories, rotormast(s), a compressor, or any other type of load or combination thereof.The compound engine assembly 10 further includes a compressor 18, aturbine section 20 compounding power with the engine core 12 and indriving engagement with the compressor 18 and generally including afirst stage turbine 22 and a second stage turbine 24, and a flowdistribution assembly 25, 125, 225, examples of which will be describedfurther below.

In a particular embodiment, the engine core 12 includes one or morerotary engine(s) drivingly engaged to the common shaft 16 driving theload and each having a rotor sealingly engaged in a respective housing,with each rotary type engine having a near constant volume combustionphase for high cycle efficiency. The rotary engine(s) may be Wankelengine(s). Referring to FIG. 2, an exemplary embodiment of a Wankelengine is shown. Each Wankel engine comprises a housing 32 defining aninternal cavity with a profile defining two lobes, which is preferablyan epitrochoid. A rotor 34 is received within the internal cavity. Therotor defines three circumferentially-spaced apex portions 36, and agenerally triangular profile with outwardly arched sides. The apexportions 36 are in sealing engagement with the inner surface of aperipheral wall 38 of the housing 32 to form three working chambers 40between the rotor 34 and the housing 32.

The rotor 34 is engaged to an eccentric portion 42 of the shaft 16 toperform orbital revolutions within the internal cavity. The shaft 16performs three rotations for each orbital revolution of the rotor 34.The geometrical axis 44 of the rotor 34 is offset from and parallel tothe axis 46 of the housing 32. During each orbital revolution, eachchamber 40 varies in volume and moves around the internal cavity toundergo the four phases of intake, compression, expansion and exhaust.

An intake port 48 is provided through the peripheral wall 38 forsuccessively admitting compressed air into each working chamber 40. Anexhaust port 50 is also provided through the peripheral wall 38 forsuccessively discharging the exhaust gases from each working chamber 40.Passages 52 for a glow plug, spark plug or other ignition element, aswell as for one or more fuel injectors (not shown) are also providedthrough the peripheral wall 38. Alternately, the intake port 48, theexhaust port 50 and/or the passages 52 may be provided through an end orside wall 54 of the housing; and/or, the ignition element and a pilotfuel injector may communicate with a pilot subchamber (not shown)defined in the housing 32 and communicating with the internal cavity forproviding a pilot injection. The pilot subchamber may be for exampledefined in an insert (not shown) received in the peripheral wall 38.

In a particular embodiment the fuel injectors are common rail fuelinjectors, and communicate with a source of Heavy fuel (e.g. diesel,kerosene (jet fuel), equivalent biofuel), and deliver the heavy fuelinto the engine(s) such that the combustion chamber is stratified with arich fuel-air mixture near the ignition source and a leaner mixtureelsewhere.

For efficient operation the working chambers 40 are sealed, for exampleby spring-loaded apex seals 56 extending from the rotor 34 to engage theperipheral wall 38, and spring-loaded face or gas seals 58 and end orcorner seals 60 extending from the rotor 34 to engage the end walls 54.The rotor 34 also includes at least one spring-loaded oil seal ring 62biased against the end wall 54 around the bearing for the rotor 34 onthe shaft eccentric portion 42.

Each Wankel engine provides an exhaust flow in the form of a relativelylong exhaust pulse; for example, in a particular embodiment, each Wankelengine has one explosion per 360° of rotation of the shaft, with theexhaust port remaining open for about 270° of that rotation, thusproviding for a pulse duty cycle of about 75%. By contrast, a piston ofa reciprocating 4-stroke piston engine typically has one explosion per720° of rotation of the shaft with the exhaust port remaining open forabout 180° of that rotation, thus providing a pulse duty cycle of 25%.

In a particular embodiment which may be particularly but not exclusivelysuitable for low altitude, each Wankel engine has a volumetric expansionratio of from 5 to 9, and a volumetric compression ratio lower than thevolumetric expansion ratio. The power recovery of the first stageturbine may be maximized by having the exhaust gas temperatures at thematerial limit, and as such is suitable for such relatively lowvolumetric compression ratios, which may help increase the power densityof the Wankel engine and may also improve combustion at high speed andof heavy fuel.

It is understood that other configurations are possible for the enginecore 12. The configuration of the engine(s) of the engine core 12, e.g.placement of ports, number and placement of seals, etc., may vary fromthat of the embodiment shown. In addition, it is understood that eachengine of the engine core 12 may be any other type of internalcombustion engine including, but not limited to, any other type ofrotary engine, and any other type of non-rotary internal combustionengine such as a reciprocating engine.

Referring back to FIG. 1, the compressor 18 is a supercharger compressorwhich may be a single-stage device or a multiple-stage device and may bea centrifugal or axial device with one or more rotors having radial,axial or mixed flow blades. Air enters the compressor and is compressedand delivered to an outlet conduit 70 communicating with the outlet 18 oof the compressor 18, and then circulated in part to an inlet conduit 71communicating with the outlet conduit 70 through the flow distributionassembly 25, 125, 225. The inlet conduit 71 delivers the compressed airto the inlet 12 i of the engine core 12, which corresponds to orcommunicates with the inlet of each engine of the engine core 12. In aparticular embodiment, the flow and pressure ratio of the compressor 18is regulated using variable inlet guide vanes (VIGV) and/or a variablediffuser at the inlet of the compressor 18 and both generally indicatedat 72, to achieve flow and power modulation. In a particular embodiment,the compressor 18 has a compression pressure ratio of approximately 4:1.Other values are also possible.

In the embodiment shown, the compressor outlet 18 o is also in fluidcommunication with a bleed conduit 74 through the flow distributionassembly 25, 125, 225, which provides a fluid communication between theoutlet conduit 70 and the bleed conduit 74. The bleed conduit 74 has anend configured for connection to a pneumatic system of the aircraft suchthat part of the compressed air from the compressor 18 may also besupplied to the aircraft to support the aircraft pneumatic system.Accordingly, the compressor 18 provides both bleed air to the aircraftand compressed air to the engine core 12.

The engine core 12 receives the pressurized air from the compressor 18and burns fuel at high pressure to provide energy. Mechanical powerproduced by the engine core 12 drives the electrical generator 14 whichprovides power for the aircraft; in the embodiment shown the connectionbetween the shaft 16 of the engine core 12 and the generator 14 is donethrough an appropriate type of gearbox 30. In another embodiment, theelectrical generator 14 has a design speed compatible with therotational speed of the engine core 12, for example from about 6000 toabout 10000 rpm (rotations per minute) with an engine core 12 includingrotary engine(s), and the shaft 16 of the engine core 12 drives theelectrical generator 14 directly (see FIG. 9)—i.e. through any type ofengagement between the engine shaft 16 with the shaft of the generatorrotor resulting in both shafts rotating at a same speed. In a particularembodiment, direct driving of the electrical generator 14 may providefor a reduction in gear losses which can be around 1% of the appliedload; in a particular embodiment, the applied load of the generator 14is about 200 hp and accordingly a loss reduction of approximately 2 hpin the waste heat produced by the engine assembly 10 may be obtained.

In a particular embodiment, the engine core 12 includes rotaryengine(s), for example Wankel engine(s), and the generator 14 directlydriven by the engine core has a nominal frequency of 400 Hz (e.g. actualfrequency range of approximately 380-420 Hz) and is a 6 pole, 3 phases,alternative current generator having a design speed of from 7600 to 8400rpm. In another particular embodiment, the generator 14 directly drivenby the rotary (e.g. Wankel) engine core has a nominal frequency of 400Hz and is a 8 pole, 3 phases, alternative current generator having adesign speed of from 5700 to 6300 rpm. In another particular embodiment,the generator 14 directly driven by the rotary (e.g. Wankel) engine corehas a nominal frequency of 400 Hz and is a 4 pole, 3 phases, alternativecurrent generator having a design speed of from 11400 to 12600 rpm.

Is it understood that other types of generators 14 may be used. Forexample, the engine assembly 10 used as an APU can be configured toprovide other high frequency alternative current supplies by selectingthe operating speed and generator pole count to provide minimum weight,volume and/or heat as required. Variable speed operation may be employedwhen the associated electrical load is not frequency sensitive. Othervariations are also possible.

The shaft 16 of the engine core 12 is also mechanically coupled to therotor(s) of the compressor 18 such as to provide mechanical powerthereto, through another gearbox 31. In a particular embodiment, thegearbox 31 providing the mechanical coupling between the rotor(s) of thecompressor 18 and the engine core 12 defines a speed ratio of about 10:1between the compressor rotor(s) and engine core.

In a particular embodiment where the engine core 12 includes internalcombustion engine(s), each engine of the engine core 12 provides anexhaust flow in the form of exhaust pulses of high pressure hot gasexiting at high peak velocity. The outlet 12 o of the engine core 12(i.e. the outlet of each engine of the engine core 12) is in fluidcommunication with the inlet 22 i of the first stage turbine 22, andaccordingly the exhaust flow from the engine core 12 is supplied to thefirst stage turbine 22. Mechanical energy recovered by the first stageturbine 22 is coupled to the shaft 16 of the engine core 12 via agearbox 33; the rotor(s) of the compressor 18 are thus drivingly engagedto the rotor(s) of the first stage turbine 22 through the engine core12. In a particular embodiment, the first stage turbine 22 is configuredas a velocity turbine, also known as an impulse turbine, and recoversthe kinetic energy of the core exhaust gas while creating minimal or noback pressure. The first stage turbine 22 may be a centrifugal or axialdevice with one or more rotors having radial, axial or mixed flowblades.

The inlet 24 i of the second stage turbine 24 is in fluid communicationwith the outlet 22 o of the first stage turbine 22 and completes therecovery of available mechanical energy from the exhaust gas. The secondturbine 24 is also coupled to the shaft 16 of the engine core 12 throughthe gearbox 33; the rotor(s) of the compressor 18 are thus drivinglyengaged to the rotor(s) of the second stage turbine 24 through theengine core 12. In a particular embodiment, the second stage turbine 24is configured as a pressure turbine, also known as a reaction turbine.The second stage turbine 24 may be a centrifugal or axial device withone or more rotors having radial, axial or mixed flow blades.

In the embodiment shown, the rotors of the first and second stageturbines 22, 24 are connected to a same shaft 23 which is coupled to theengine core 12 through the gearbox 33. Alternately, the turbines 22, 24could be mounted on different shafts, for example with the first stageturbine 22 mounted on a first shaft coupled to the engine shaft 16 (forexample through the gearbox 23) and the second stage turbine 24 mountedon a second shaft drivingly engaged to the compressor 18.

A pure impulse turbine works by changing the direction of the flowwithout accelerating the flow inside the rotor; the fluid is deflectedwithout a significant pressure drop across the rotor blades. The bladesof the pure impulse turbine are designed such that in a transverse planeperpendicular to the direction of flow, the area defined between theblades is the same at the leading edges of the blades and at thetrailing edges of the blade: the flow area of the turbine is constant,and the blades are usually symmetrical about the plane of the rotatingdisc. The work of the pure impulse turbine is due only to the change ofdirection in the flow through the turbine blades. Typical pure impulseturbines include steam and hydraulic turbines.

In contrast, a reaction turbine accelerates the flow inside the rotorbut needs a static pressure drop across the rotor to enable this flowacceleration. The blades of the reaction turbine are designed such thatin a transverse plane perpendicular to the direction of flow, the areadefined between the blades is larger at the leading edges of the bladesthan at the trailing edges of the blade: the flow area of the turbinereduces along the direction of flow, and the blades are usually notsymmetrical about the plane of the rotating disc. The work of the purereaction turbine is due mostly to the acceleration of the flow throughthe turbine blades.

Most aeronautical turbines are not “pure impulse” or “pure reaction”,but rather operate following a mix of these two opposite butcomplementary principles—i.e. there is a pressure drop across theblades, there is some reduction of flow area of the turbine blades alongthe direction of flow, and the speed of rotation of the turbine is dueto both the acceleration and the change of direction of the flow. Thedegree of reaction of a turbine can be determined using thetemperature-based reaction ratio (equation 1) or the pressure-basedreaction ratio (equation 2), which are typically close to one another invalue for a same turbine:

$\begin{matrix}{{{Reaction}(T)} = \frac{( {t_{S\; 3} - t_{S\; 5}} )}{( {t_{S\; 0} - t_{S\; 5}} )}} & (1) \\{{{Reaction}(P)} = \frac{( {P_{S\; 3} - P_{S\; 5}} )}{( {P_{S\; 0} - P_{S\; 5}} )}} & (2)\end{matrix}$where T is temperature and P is pressure, s refers to a static port, andthe numbers refers to the location the temperature or pressure ismeasured: 0 for the inlet of the turbine vane (stator), 3 for the inletof the turbine blade (rotor) and 5 for the exit of the turbine blade(rotor); and where a pure impulse turbine would have a ratio of 0 (0%)and a pure reaction turbine would have a ratio of 1 (100%).

In a particular embodiment, the first stage turbine 22 is configured totake benefit of the kinetic energy of the pulsating flow exiting theengine core 12 while stabilizing the flow, and the second stage turbine24 is configured to extract energy from the remaining pressure in theflow while expanding the flow. Accordingly, the first stage turbine 22has a smaller reaction ratio than that of the second stage turbine 24.

In a particular embodiment, the second stage turbine 24 has a reactionratio higher than 0.25; in another particular embodiment, the secondstage turbine 24 has a reaction ratio higher than 0.3; in anotherparticular embodiment, the second stage turbine 24 has a reaction ratioof about 0.5; in another particular embodiment, the second stage turbine24 has a reaction ratio higher than 0.5.

In a particular embodiment, the first stage turbine 22 has a reactionratio of at most 0.2; in another particular embodiment, the first stageturbine 22 has a reaction ratio of at most 0.15; in another particularembodiment, the first stage turbine 22 has a reaction ratio of at most0.1; in another particular embodiment, the first stage turbine 22 has areaction ratio of at most 0.05.

It is understood that any of the above-mentioned reaction ratios for thesecond stage turbine 24 can be combined with any of the above-mentionedreaction ratios for the first stage turbine 22, and that these valuescan correspond to pressure-based or temperature-based ratios. Othervalues are also possible. For example, in a particular embodiment, thetwo turbines 22, 24 may have a same or similar reaction ratio; inanother embodiment, the first stage turbine 22 has a higher reactionratio than that of the second stage turbine 24. Both turbines 22, 24 maybe configured as impulse turbines, or both turbines 22, 24 may beconfigured as pressure turbines.

Is it understood that the connections between the rotors of thecompressor 18 and turbines 22, 24 may be different than the embodimentshown. For example, the rotors of the compressor 18 and turbines 22, 24may be coupled to the engine core 12 by a gearing system or variablespeed drive such that power can be shared mechanically. Alternately, thecompressor may be a turbocharger directly driven by the second stageturbine 24 without power transfer between the compressor 18 and theengine core 12, for example by having the rotors of the compressor 18and second stage turbine 24 mounted on common shaft rotatingindependently of the shaft 16 of the engine core 12. In this case avariable area turbine vane may be provided at the inlet of the secondstage (turbocharger) turbine 24 to provide adequate control of thecompressor drive.

In use, there are typically operational situations where the aircraftcannot accept compressed air from the APU but still requires the APU torun, for example to power the generator 14. In this case the compressor18 produces excess flow and needs to be protected from surge. In theembodiment shown, the compressor outlet 18 o is also in fluidcommunication with an excess air duct 82 receiving this excess or surgeflow, through the flow distribution assembly 25, 125, 225 which providesa fluid communication between the outlet conduit 70 and the excess airduct 82. The excess air duct 82 provides an alternate path for theexcess air produced by the compressor 18.

In a particular embodiment which is not shown, the excess air is dumpedto atmosphere, for example by having the excess air duct 82 in fluidcommunication with the exhaust of the engine assembly 10. In theembodiment shown, the excess air duct 82 has a first end communicatingwith the compressor outlet 18 o and an opposed end communicating withthe second stage turbine inlet 24 i, such as to recover energy from themain flow and the surge excess flow. The excess air duct 82 thus definesa flow path between the compressor outlet 18 o and the turbine sectionwhich is separate from the engine core 12. The excess air duct 82 maycommunicate with the second stage turbine inlet 24 i together with theexhaust from the first stage turbine outlet 22 o through an inlet mixingdevice, or through a partial segregated admission turbine configuration(segregated admission nozzle) where some vane passages in the turbineentry nozzle are dedicated to the flow from the excess air duct 82 whileother vane passages are dedicated to the exhaust flow from the firststage turbine outlet 22 o. The second stage turbine 24 may feature avariable nozzle to facilitate control of load sharing and differentlevels of returned excess air.

The excess air duct 82 may alternately communicate with the inlet 22 iof the first stage turbine 22, or with the inlet of a third turbine (notshown) dedicated to recovering excess air energy. Such a third turbinemay be connected to the shaft 16 of the engine core 12, for examplethrough an over-running clutch, to return the energy extracted from theexcess flow to the shaft 16, or may be used to drive other elements,including, but not limited to, a cooling fan and/or an additionalgenerator. Other types of connections and configurations are alsopossible.

In a particular embodiment, an exhaust heat exchanger 28 is provided toprovide heat exchange relationship between the air circulating throughthe excess air duct 82 and the exhaust air from the outlet 24 o of thesecond stage turbine 24. The heat exchanger 28 thus includes at leastone first conduit 28 a in heat exchange relationship with at least onesecond conduit 28 b. The excess air duct 82 is in fluid communicationwith the second stage turbine inlet 24 i through the first conduit(s) 28a of the heat exchanger 28, and the second conduit(s) 28 b of the heatexchanger 28 is/are in fluid communication with the second stage turbineoutlet 24 o such that the exhaust from the second stage turbine 24circulates therethrough. In a particular embodiment, the exhaust heatexchanger 28 recovers energy from the waste heat in the exhaust andincreases the temperature of the excess flow (surge bleed flow) cominginto the second stage turbine 24, which improves its capacity to do workin the turbine. This provides a hybrid partially recuperated cycle.

In an alternate embodiment, the exhaust heat exchanger 28 is omitted.

Exemplary embodiments for the flow diverting assembly 25, 125, 225 willnow be described. It is understood however than the compressor outlet 18o/outlet conduit 70 can be in fluid communication with the engine coreinlet 12 i/inlet conduit 71, the bleed conduit 74 and/or the excess airduct 82 through any other appropriate type or configuration of fluidcommunication. For example, the bleed conduit 74 could be connected tothe compressor outlet 18 o separately from the outlet conduit 70.

Referring to FIG. 3, in a particular embodiment, the flow divertingassembly 25 includes an intercooler 26, and the outlet conduit 70 isconnected to a branch 73 splitting the flow between a bypass conduit 78,an intercooler inlet conduit 64 and the excess air duct 82. Thecommunication between the branch 73 and the excess air duct 82 isperformed through a diverter valve 84 to effect throttling of the excessair flow or shut it off when required. The excess air duct 82 thuscommunicates with the outlet conduit 70 upstream of the intercooler 26;in a particular embodiment, such a configuration allows for leavingmaximum energy in the compressed air being diverted into the excess airduct 82.

The intercooler 26 includes at least one first conduit 26 a in heatexchange relationship with at least one second conduit 26 b. Each firstconduit 26 a of the intercooler 26 has an inlet in fluid communicationwith the intercooler inlet conduit 64, and an outlet in fluidcommunication with an intercooler outlet conduit 66. Each second conduit26 b of the intercooler 26 is configured for circulation of a coolanttherethrough, for example cooling air. The compressed air circulatingthrough the first conduit(s) 26 a is thus cooled by the coolantcirculating through the second conduit(s) 26 b.

The intercooler outlet conduit 66 is in fluid communication with theinlet conduit 71 (and accordingly with the engine core inlet 12 i) andwith the bleed conduit 74; the communication with the bleed conduit 74is performed through a bleed air valve 76, which in a particularembodiment is a load control valve, to effect throttling of the bleed orshut it off when required. The intercooler 26 accordingly reduces thetemperature of the compressed air going to the engine core 12 as well asthe compressed air being channelled to the aircraft through the bleedair valve 76 and bleed conduit 74. In a particular embodiment, thepre-cooling of the air going to the aircraft allows for a higherdelivery pressure than APU systems which are not pre-cooled andaccordingly temperature limited for safety reasons. A higher pressuredelivery may generally permit smaller ducts and pneumatic equipment,which may allow for weight savings on the aircraft.

The bypass conduit 78 provides fluid communication between the outletconduit 70 and each of the inlet conduit 71 and bleed air valve 76 inparallel of the intercooler 26, thus allowing for a selected part of theflow to bypass the intercooler 26 before reaching the bleed air valve 76(and accordingly the bleed conduit 74) and the inlet conduit 71 (andaccordingly the engine core inlet 12 i). The bypass conduit 78 includesa bypass valve 80 regulating the flow bypassing the intercooler 26.Accordingly, the temperature of the compressed air circulated to theinlet and bleed conduits 71, 74 may be regulated by changing theproportion of the flow going through the intercooler 26 by controllingthe proportion of the flow going through the bypass conduit 78 with thebypass valve 80. In this particular embodiment, the compressed flowcirculated to the inlet conduit 71 has the same temperature as thecompressed flow circulated to the bleed conduit 74. In a particularembodiment the compressed air is cooled by the intercooler 26 such thatthe air circulated to the bleed conduit 74 and the inlet conduit 71 hasa temperature of 250° F. or lower; other values are also possible.

Referring to FIG. 4, another particular embodiment of the flow divertingassembly 125 is shown. The outlet conduit 70 is connected to the bypassconduit 78, the intercooler inlet conduit 64, the excess air duct 82(through the diverter valve 84) and the bleed conduit 74 (through thebleed air valve 76). In this embodiment, since the bleed conduit 74communicates with the outlet conduit 70 upstream of the intercooler 26,the compressed air is not cooled before being circulated to the bleedconduit 74.

The intercooler outlet conduit 66 is in fluid communication with theinlet conduit 71 (and accordingly with the engine core inlet 12 i); theintercooler 26 thus reduces the temperature of the compressed air goingto the engine core 12. The bypass conduit 78 provides fluidcommunication between the outlet conduit 70 and the inlet conduit 71 inparallel of the intercooler 26, thus allowing for a selected part of theflow to bypass the intercooler 26 before reaching the inlet conduit 71(and accordingly the engine core inlet 12 i). The temperature of thecompressed air circulated to the inlet conduit 71 may be regulated bychanging the proportion of the flow going through the intercooler 26 bycontrolling the proportion of the flow going through the bypass conduit78 with the bypass valve 80 included therein. In a particular embodimentthe compressed air circulating in the outlet conduit 70 and to the bleedconduit 74 has a temperature of 450° F. or lower, and the intercooler 26cools part of the compressed air so that the air circulated to the inletconduit 71 has a temperature of 250° F. or lower; other values are alsopossible. In a particular embodiment, using the intercooler 26 to coolonly the portion of the compressed air circulated to the inlet conduit71 may allow for the intercooler 26 to be significantly smaller than anintercooler also used to cool the portion of the air circulated to thebleed conduit 74, for example such as shown in FIG. 3.

Referring to FIG. 5, another particular embodiment of the flow divertingassembly 225 is shown. The outlet conduit 70 is connected to the bypassconduit 78, the intercooler inlet conduit 64, and the excess air duct 82(through the diverter valve 84). A first intercooler 126 used as apre-cooler has first conduit(s) 126 a each having an inlet in fluidcommunication with the intercooler inlet conduit 64, and an outlet influid communication with an intercooler intermediate conduit 68.

The intercooler intermediate conduit 68 is in fluid communication withthe bleed conduit 74 through the bleed air valve 76, and the bypassconduit 78 provides fluid communication between the outlet conduit 70and a portion of the bleed conduit 74 upstream of the bleed air valve 76in parallel of the intercooler 126. Accordingly, the temperature of thecompressed air circulated to the bleed conduit 74 may be regulated bychanging the proportion of the flow going through the intercooler 126 bycontrolling the proportion of the flow going through the bypass conduit78 with the bypass valve 80.

The flow diverting assembly 225 includes a second intercooler 226 alsohaving first conduit(s) 226 a in heat exchange relationship with secondconduit(s) 226 b. Each first conduit 226 a of the intercooler 226 has aninlet in fluid communication with the intercooler intermediate conduit68, and an outlet in fluid communication with the intercooler outletconduit 66. Each second conduit 226 b of the intercooler 226 isconfigured for circulation of a coolant therethrough, for examplecooling air.

The intercooler outlet conduit 66 is in fluid communication with theinlet conduit 71 (and accordingly with the engine core inlet 12 i); thesecond intercooler 226 thus further reduces the temperature of thecompressed air going to the engine core 12. An additional bypass conduit178 provides fluid communication between the intercooler intermediateconduit 68 and the inlet conduit 71 in parallel of the intercooler 226,thus allowing for a selected part of the flow to bypass the intercooler226 before reaching the inlet conduit 71 (and accordingly the enginecore inlet 12 i). The additional bypass conduit 178 includes anadditional bypass valve 180 to regulate the flow circulatingtherethrough. The temperature of the compressed air circulated to theinlet conduit 71 may be regulated by changing the proportion of the flowgoing through the intercoolers 126, 226 by controlling the proportion ofthe flow going through the bypass conduits 78, 178 with the bypassvalves 80, 180.

An alternate embodiment is shown in dotted lines, where the bypassconduit 178 is replaced by a bypass conduit 178′ containing bypass valve180′ and extending between the outlet conduit 70 and inlet conduit 71.

The inlet conduit 71 thus communicates with the pre-cooler intercooler126 at least in part through the second intercooler 226, while the bleedconduit 74 communicates with the pre-cooler intercooler 126 upstream ofthe second intercooler 226, thus independently thereof. The arrangementthus allows for separate regulation of the temperature of the flowreaching the bleed conduit 74 and of the flow reaching the inlet conduit71. For example, the proportion of the flow circulating through theintercooler 126 may be selected such that the temperature of the flowreaching the bleed conduit 74 is 450° F. or lower, and the temperatureof the flow is further reduced in the second intercooler 226 to have avalue of 250° F. or lower when reaching the inlet conduit 71. Othervalues are also possible.

In all embodiments, the flow diverting assembly 25, 125, 225 may includepressure, temperature and/or flow sensors, and/or closed loop system(s)controlling the position of one, some or all of the valves 76, 80, 84,180, 180′. Any one, some or all of the valves 76, 80, 84, 180, 180′ maybe a hydraulically, pneumatically or electrically driven modulatingvalve.

In a particular embodiment, the engine assembly 10 is air startable fromthe pneumatic system of the aircraft or the engine bleed, as opposed toelectrical power. Opening the bleed air valve 76 and the diverter valve84 admits pressurized air to the second stage turbine 24, therebyproviding a means to start rotation of the engine assembly 10. Such aconfiguration may thus allow for a rapid start in flight without theneed to use electrical power. When provided, the third turbine (notshown) receiving air from the excess air duct 82 could also allow forair starting of the engine assembly 10. Check valves or bypass valves(not shown) may be needed to prevent reverse flow through other parts ofthe engine assembly 10. In both cases the compressor 18 is dead headedon the bleed side, so the engine start at low speed, and the externalstart flow is cancelled as soon as possible before accelerating to fullspeed to prevent high energy compressor stalls.

In an alternate embodiment, the diverter valve 84 is omitted or may besimplified to a two position on/off valve. The flow from the compressor18 is ducted to the engine core 12 and to the excess air duct 82, andthe compressed air is “bled” from the excess air duct 82 as required bythe aircraft, limited if necessary by the load control valve 76. Such aconfiguration may allow loss reduction by eliminating the regulating ordiverter valve pressure drop between the compressor 18 and thedownstream heat exchanger 28 and turbine. When a two position divertervalve 84 is employed the valve is closed when the aircraft has a highpneumatic demand on the engine assembly 10 and fully open when theengine assembly 10 is operated with low pneumatic demand or forelectrical power only.

In an alternate embodiment, the intercoolers 26, 126, 226 are omittedand the flow is circulated to the inlet conduit 71 and bleed conduit 74without being cooled.

Referring to FIG. 6, in a particular embodiment, the engine assembly 10includes an oil cooler 88 to remove heat from the oil system of theengine assembly 10, and an engine core liquid cooler 89 to remove heatfrom the coolant (e.g. water, oil or other liquid coolant) of thecooling system of the engine core 12. A coolant pump 94 circulates thecoolant between the engine core 12 and the engine core liquid cooler 89.The coolers 88, 89 are integrated with the intercooler 26/226 (andpre-cooler intercooler 126 if provided) in a cooling assembly to preventreplication of items like cooling fans and eductors. In this particularembodiment, the coolers 88, 89 and intercoolers 26/226, 126 are arrangedin series in a single air duct 90 which is vented by a cooling fan 92 inground operations or part of a ram air circuit in flight. The coolingfan 92 may be driven by any suitable rotating element of the engineassembly 10, or powered by the generator 14. A single inlet and exhaustcan thus be used for providing coolant to all the coolers 88, 89, andintercoolers 26/226, 126. The coolers 88, 89 and intercooler 26/226, 126are placed within the duct according to their temperature requirements.In the embodiment shown, the oil cooler 88 and engine core liquid cooler89 have the lowest temperature requirements (e.g. requiring to cool thefluids therein at around 180° F. to 200° F.), and the cooling airtemperature at the inlet of the air duct 90 is 130° F. or lower; theintercooler 26/226 has a higher temperature requirement than the coolers88, 89 (e.g. around 250° F.), and the pre-cooler intercooler 126 (ifprovided) has a higher temperature requirement than the intercooler26/226 (e.g. around 450° F.). Other values are also possible.

Referring to FIG. 7, another embodiment for the cooling assembly isshown. In this embodiment, the coolers 88, 89 and the intercooler 26/226are in parallel in the air duct 190 vented by the cooling fan 92, withthe pre-cooler intercooler 126 being provided downstream of the others.In another embodiment which is not shown, the coolers 88, 89 andintercoolers 26/226, 126 are all placed in parallel in the air duct.

Referring to FIG. 8, a compound engine assembly 110 with coolingassembly in accordance with a particular embodiment is shown. In thisembodiment, the second conduit(s) 226 b of the intercooler 226 are influid communication with the liquid cooling system of the engine core 20such that the coolant from the liquid cooling system is circulated inthe second conduit(s) 226 b to cool the compressed air circulated in thefirst conduit(s) 226 a. In this embodiment, a mechanical drive 96 isprovided between the cooling fan 92 and the shaft 16 of the engine core12. The coolers 88, 89 are placed in parallel in the cooling air duct290 upstream of the fan 92, and the pre-cooler intercooler 126 is placedin the duct 290 downstream of the fan 92. In a particular embodiment,such an arrangement provides for an optimal cooling delta T to the oil,engine coolant and compressed air while preserving acceptable entrytemperatures to the fan 92 to keep its power below a desirable thresholdand avoid the need for more expensive materials which may be required ifthe fan 92 was located in a hotter zone downstream of the intercooler126.

Referring to FIG. 9, a compound engine assembly 210 according to anotherembodiment is shown, where components similar to that of the embodimentshown in FIG. 1 are identified by the same reference numerals and arenot further described herein. As described above, the engine core 12includes one or more internal combustion engine(s) including, but notlimited to, any type of rotary engine (e.g. Wankel engine), and any typeof non-rotary internal combustion engine such as a reciprocating engine.Any one of the flow distribution assemblies 25, 125, 225 or any otherappropriate flow distribution assembly may be used to distribute theflow from the outlet conduit 70 to the inlet conduit 71, bleed conduit74 and excess air duct 84.

In this embodiment, the shaft 16 of the engine core 12 is onlymechanically coupled to the generator 14, and not to the rotors of thecompressor 18 and turbines 22, 24. The first and second stage turbines22, 24 are mechanically coupled to the compressor 18, for example byhaving their rotors supported by a same turbine shaft 123. The first andsecond stage turbines 22, 24 are also mechanically coupled to a secondgenerator/electrical motor 114. In a particular embodiment, the shaft 16of the engine core 12 and the turbine shaft 123 are each coupled totheir respective generator 14, 114 through a direct connection; thegenerator 14 coupled to the engine core 12 may thus have a lower speedof rotation than the generator 114 coupled to the turbine shaft 123.Alternately, one or both connection(s) may be performed through arespective gearbox (not shown).

In another particular embodiment, the generator 114 directly driven bythe turbines 22, 24 (or by one of the turbines in an embodiment wherethe turbines are on different shafts) has a nominal frequency of atleast 400 Hz suitable for high power density aircraft electricalequipment, for example a 2 poles, alternative current generator with anominal frequency of 400 Hz having a design speed of from 22800 to 25300rpm, which may correspond to a nominal speed of 24000 rpm. Such agenerator 114 may be used in combination with any of the particulargenerators 14 mentioned above.

Power from the two shafts 16, 123 is compounded through electrical powerbeing transferred between the two generators 14, 114. For example, thefirst generator 14 transfers power to the second generator/motor 114which acts as a motor to drive the rotor(s) of the compressor 18. Thegenerators 14, 114 also provide electrical power for the aircraft.

In the embodiment shown, a power controller 86 is provided to controlpower transfer between the two generators 14, 114 and power provided tothe aircraft. In a particular embodiment, the power controller 86 allowsfor the compressor and engine core speed ratio to be variable, with eachrotational speed being scheduled independently for optimal performance.The portion of the power from the first generator 14 being transferredto the second generator/motor 114 can be controlled to achieve the mostadvantageous rotational speed for the rotor(s) of the compressor 18. Inaddition, when the turbines 22, 24 coupled to the compressor 18 generateexcess energy, the second generator/motor 114 can also provide power tothe aircraft and/or to the first generator 14 which may also act as amotor. The power controller 86 may also contain features such asfrequency and voltage regulation to manage AC power quality supplied tothe airframe.

Although not shown, the transfer of power from the engine core 12 to thesupercharger (compressor 18 and turbines 22, 24) may also be performedthrough hydraulic or mechanical CVT systems to allow for independentspeed scheduling.

Referring to FIG. 10, an engine assembly 310 according to anotherembodiment is shown, where components similar to that of the embodimentshown in FIG. 1 are identified by the same reference numerals and arenot further described herein.

In this embodiment, the shaft 16 of the engine core 12 is mechanicallycoupled to the generator 14, and the power of the turbine(s) is notcompounded with that of the engine core 12. Although a single turbine322 is shown, multiple turbines may be provided. The turbine 322 ismechanically coupled to the compressor 18, for example by having theirrotors supported by a same shaft 123. The turbine 322 may be configuredas an impulse turbine or as a pressure turbine, and may have anysuitable reaction ratio, including, but not limited to, the ratiosdescribed above for turbines 22, 24. In a particular embodiment, theturbine 322 is replaced by first and second stages turbines 22, 24 aspreviously described.

Any one of the flow distribution assemblies 25, 125, 225 or any otherappropriate flow distribution assembly may be used to distribute theflow from the outlet conduit 70 to the inlet conduit 71, bleed conduit74 and excess air duct 84. When more than one turbine is provided, theflow from the excess air duct 84 may be circulated to the inlet of anyone of the turbines.

Although not shown, the turbine(s) 322 may also drive a separategenerator or any other appropriate type of accessory. Although notshown, a power controller may be provided to control power transferbetween the generator 14 and any system receiving electrical power fromthe generator 14.

The above description is meant to be exemplary only, and one skilled inthe art will recognize that changes may be made to the embodimentsdescribed without departing from the scope of the invention disclosed.For example, although the compressor 18 has been shown as providingcompressed air both for the engine core 12 and for the aircraft,alternately the compressor 18 may be configured to act only as asupercharger for the engine core 12, and a separate load compressor maybe configured to provide the aircraft air. Such a load compressor may bedriven by the engine core 12 and/or the turbines 22, 24, 322 eitherdirectly or through a gearbox. The two compressors may have a commoninlet. Moreover, although the engine core 12 has been described asincluding one or more internal combustion engines, the engine core 12may alternately be any other type of engine core in which the compressedair is mixed with fuel and ignited for generating hot combustion gases,including, but not limited to, a gas turbine engine combustor; asnon-limiting examples, the intercooler 26, 126 cooling compressed aircirculated to the aircraft, and the circulation of the excess air fromthe compressor 18 to a turbine 22, 24 with the excess air duct 82 toprovide additional work with or without an exhaust heat exchanger 28between the excess air and the turbine exhaust, may be applied to a gasturbine engine with a combustor, as schematically illustrated in FIG.11. Other modifications which fall within the scope of the presentinvention will be apparent to those skilled in the art, in light of areview of this disclosure, and such modifications are intended to fallwithin the appended claims.

The invention claimed is:
 1. A power unit for an aircraft, the powerunit comprising: at least one internal combustion engine in drivingengagement with an engine shaft; a generator having a generator shaft indriving engagement with the engine shaft to provide electrical power forthe aircraft, the generator and engine shafts being connected via agearbox such as to be rotatable at a different speed; a compressorhaving an outlet in fluid communication with an inlet of the at leastone internal combustion engine; and a turbine having an inlet in fluidcommunication with an outlet of the at least one internal combustionengine.
 2. The power unit as defined in claim 1, wherein the turbine isconfigured to compound power with the at least one internal combustionengine.
 3. The power unit as defined in claim 1, further comprising ableed conduit having an end configured for connection with a pneumaticsystem of the aircraft, the bleed conduit in fluid communication withthe outlet of the compressor through a bleed air valve selectivelyopening and closing the fluid communication between the outlet of thecompressor and the end of the bleed conduit configured for connection tothe pneumatic system.
 4. The power unit as defined in claim 1, whereineach of the at least one internal combustion engine includes a rotorsealingly and rotationally received within a respective internal cavityto provide rotating chambers of variable volume in the respectiveinternal cavity, the rotor having three apex portions separating therotating chambers and mounted for eccentric revolutions within therespective internal cavity, the respective internal cavity having anepitrochoid shape with two lobes.
 5. The power unit as defined in claim1, wherein the generator has a nominal rotational speed of at least 5700rotations per minute.
 6. The power unit as defined in claim 1, whereinthe generator is selected from the group consisting of: a 6 pole, 3phases, alternative current generator having a design speed of from 7600to 8400 rotations per minute; a 8 pole, 3 phases, alternative currentgenerator having a design speed of from 5700 to 6300 rotations perminute; and a 4 pole, 3 phases, alternative current generator having adesign speed of from 11400 to 12600 rotations per minute.
 7. The powerunit as defined in claim 1, wherein the generator is a first generator,the turbine includes at least one rotor engaged on a turbine shaftrotatable independently of the engine shaft, the assembly furthercomprising a second generator having a second generator shaft directlyengaged to the turbine shaft such that the turbine and second generatorshafts are rotatable at a same speed.
 8. The power unit as defined inclaim 1, wherein the turbine is a first stage turbine, the assemblyfurther comprising a second stage turbine having an inlet in fluidcommunication with an outlet of the first stage turbine.
 9. The powerunit as defined in claim 8, wherein the first stage turbine isconfigured as an impulse turbine with a pressure-based reaction ratiohaving a value of at most 0.25, the second stage turbine having a higherreaction ratio than that of the first stage turbine.
 10. The power unitas defined in claim 1, further comprising variable inlet guide vanes, avariable diffuser or a combination thereof at an inlet of thecompressor.
 11. An engine assembly for use as a power unit for anaircraft, the engine assembly comprising: at least one internalcombustion engine in driving engagement with an engine shaft; agenerator having a generator shaft in driving engagement with the engineshaft to provide electrical power for the aircraft; a gearbox drivinglyconnecting the engine shaft to the generator shaft; a compressor havingan outlet in fluid communication with an inlet of the at least oneinternal combustion engine; a first stage turbine having an inlet influid communication with an outlet of the at least one internalcombustion engine; and a second stage turbine having an inlet in fluidcommunication with an outlet of the first stage turbine; wherein atleast one of the turbines is in driving engagement with the compressor.12. The engine assembly as defined in claim 11, further comprising ableed conduit having an end configured for connection to a system of theaircraft, the bleed conduit being in fluid communication with the outletof the compressor, and a bleed air valve selectively opening and closingthe fluid communication between the end of the bleed conduit and theoutlet of the compressor.
 13. The engine assembly as defined in claim11, wherein each of the at least one internal combustion engine includesa rotor sealingly and rotationally received within a respective internalcavity to provide rotating chambers of variable volume in the respectiveinternal cavity, the rotor having three apex portions separating therotating chambers and mounted for eccentric revolutions within therespective internal cavity, the respective internal cavity having anepitrochoid shape with two lobes.
 14. The engine assembly as defined inclaim 11, wherein the generator has a nominal frequency of 400 Hz. 15.The engine assembly as defined in claim 11, wherein the generator isselected from the group consisting of: a 6 pole, 3 phases, alternativecurrent generator having a design speed of from 7600 to 8400 rotationsper minute; a 8 pole, 3 phases, alternative current generator having adesign speed of from 5700 to 6300 rotations per minute; and a 4 pole, 3phases, alternative current generator having a design speed of from11400 to 12600 rotations per minute.
 16. The engine assembly as definedin claim 11, wherein rotors of the first and second stage turbines andof the compressor are engaged on a turbine shaft rotatable independentlyof the engine shaft, the generator is a first generator, the assemblyfurther comprising a second generator having a second generator shaftdirectly engaged to the turbine shaft such that the turbine and secondgenerator shafts are rotatable at a same speed.
 17. The engine assemblyas defined in claim 11, wherein the first stage turbine is configured asan impulse turbine with a pressure-based reaction ratio having a valueof at most 0.25, the second stage turbine having a reaction ratio higherthan that of the first stage turbine.
 18. A method of providingelectrical power to an aircraft, the method comprising: flowingcompressed air from an outlet of a compressor to an inlet of at leastone internal combustion engine of an engine assembly; rotating an engineshaft with the at least one internal combustion engine; driving agenerator providing electrical power to the aircraft with the engineshaft through a drive engagement between a shaft of the generator andthe engine shaft, the shaft of the generator being driven at a differentspeed than a rotational speed of the engine shaft; and driving a turbineof the engine assembly with exhaust from the at least one internalcombustion engine.